Impingement cooling arrangement for airfoils

ABSTRACT

An airfoil for a gas turbine engine according to an example of the present disclosure includes, among other things, an airfoil section that has an internal wall and an external wall. The external wall defines pressure and suction sides that extends in a chordwise direction between a leading edge and a trailing edge, a first impingement cavity and a second impingement cavity bounded by the external wall at a leading edge region that defines the leading edge. A first crossover passage within the internal wall is connected to the first impingement. The first crossover passage defines a first passage axis that intersects a surface of the first impingement cavity. A second crossover passage within the internal wall is connected to the second impingement cavity. The second crossover passage defines a second passage axis that intersects a surface of the second impingement cavity.

CROSS-REFERENCE TO RELATED APPLICATION

The present disclosure is a continuation of U.S. patent application Ser.No. 15/866,557, filed Jan. 10, 2018, herein incorporated by reference inits entirety.

BACKGROUND

This disclosure relates to cooling for a component of a gas turbineengine, including a component having one or more impingement coolingfeatures.

Gas turbine engines can include a fan for propulsion air and to coolcomponents. The fan also delivers air into an engine core where it iscompressed. The compressed air is then delivered into a combustionsection, where it is mixed with fuel and ignited. The combustion gasexpands downstream and drives turbine blades. Static vanes arepositioned adjacent to the turbine blades to control the flow of theproducts of combustion. The blades and vanes are subject to extremeheat, and thus cooling schemes are utilized for each.

SUMMARY

An airfoil for a gas turbine engine according to an example of thepresent disclosure includes an airfoil section that has an internal walland an external wall. The external wall defines pressure and suctionsides that extends in a chordwise direction between a leading edge and atrailing edge, a first impingement cavity and a second impingementcavity bounded by the external wall at a leading edge region thatdefines the leading edge, a first feeding cavity separated from thefirst impingement cavity and from the second impingement cavity by theinternal wall, and a first crossover passage within the internal wallthat connects the first impingement cavity and the first feeding cavity.The first crossover passage defines a first passage axis that intersectsa surface of the first impingement cavity. A second crossover passagewithin the internal wall connects to the second impingement cavity. Thesecond crossover passage defines a second passage axis that intersects asurface of the second impingement cavity.

In a further embodiment of any of the foregoing embodiments, the secondcrossover passage connects the second impingement cavity and a secondfeeding cavity. The second feeding cavity separates from the firstimpingement cavity and from the second impingement cavity by theinternal wall.

In a further embodiment of any of the foregoing embodiments, the firstpassage axis intersects the surface of the first impingement cavityadjacent to the pressure side, and the second passage axis intersectsthe surface of the second impingement cavity adjacent to the suctionside.

In a further embodiment of any of the foregoing embodiments, the firstcrossover passage extends between a first inlet and a first outlet. Thesecond crossover passage extends between a second inlet and a secondoutlet, and the first inlet is forward of the second outlet with respectto the chordwise direction.

In a further embodiment of any of the foregoing embodiments, the secondcrossover passage connects the second impingement cavity and the firstfeeding cavity, and the first outlet is forward of the second inlet withrespect to the chordwise direction.

In a further embodiment of any of the foregoing embodiments, the firstimpingement cavity defines a first volume and the second impingementcavity defines a second volume that is less than the first volume.

In a further embodiment of any of the foregoing embodiments, the airfoilsection extends in a spanwise direction from a 0% span position to a100% span position, and each of an outlet of the first crossover passageand an outlet of the second crossover passage is defined at a spanposition that is between an 70% span position and the 100% spanposition.

In a further embodiment of any of the foregoing embodiments, the airfoilis a high lift airfoil.

In a further embodiment of any of the foregoing embodiments, the airfoilis a turbine blade.

In a further embodiment of any of the foregoing embodiments, the airfoilsection defines a mean camber line extending between the leading andtrailing edges to bisect a thickness of the airfoil section, and a ribseparating the first impingement cavity and the second impingementcavity being transverse to the mean camber line.

In a further embodiment of any of the foregoing embodiments, the firstpassage axis intersects the external wall at an intersection point. Theintersection point is between the trailing edge and a gauge pointdefined by the airfoil.

A core for an airfoil according to an example of the present disclosureincludes a first portion that corresponds to a first impingement cavityof a leading edge region of an airfoil, a second portion thatcorresponds to a second impingement cavity of the leading edge region ofthe airfoil, a third portion that corresponds to a first feeding cavityof the airfoil, and a fourth portion that corresponds to a secondfeeding cavity of the airfoil. A first set of connectors couple thefirst portion and the third portion. The first set of connectorscorrespond to a first set of crossover passages of the airfoil. A secondset of connectors couple the second portion and the fourth portion. Thesecond set of connectors correspond to a second set of crossoverpassages of the airfoil.

In a further embodiment of any of the foregoing embodiments, the firstportion defines a first volume and the second portion defines a secondvolume that is less than the first volume.

In a further embodiment of any of the foregoing embodiments, the airfoilis a turbine airfoil.

A gas turbine engine according to an example of the present disclosureincludes an array of airfoils circumferentially distributed about anengine axis. Each airfoil of the array of airfoils has an airfoilsection that has an internal wall and an external wall. The externalwall defines pressure and suction sides that extends in a chordwisedirection between a leading edge and a trailing edge. The airfoilsection has a first impingement cavity and a second impingement cavitybounded by the external wall at a leading edge region that defines theleading edge. A first feeding cavity and a second feeding cavity areseparated from the first impingement cavity and from the secondimpingement cavity by the internal wall. A first set of crossoverpassages within the internal wall connects the first impingement cavityand the first feeding cavity. Each passage of the first set of crossoverpassages defines a first passage axis that intersects a surface of thefirst impingement cavity. A second set of crossover passages within theinternal wall connects the second impingement cavity and the secondfeeding cavity. Each passage of the second set of crossover passagesdefines a second passage axis that intersects a surface of the secondimpingement cavity.

A further embodiment of any of the foregoing embodiments includes acompressor section and a turbine section. The array of airfoils arelocated in at least one of the compressor section and the turbinesection.

In a further embodiment of any of the foregoing embodiments, the arrayof airfoils are rotatable blades.

In a further embodiment of any of the foregoing embodiments, eachpassage of the first set of crossover passages extends between a firstinlet and a first outlet, each passage of the second set of crossoverpassages extends between a second inlet and a second outlet, and boththe first inlet and the first outlet are forward of the second outletwith respect to the chordwise direction.

In a further embodiment of any of the foregoing embodiments, the firstimpingement cavity defines a first volume and the second impingementcavity defines a second volume that is less than the first volume.

In a further embodiment of any of the foregoing embodiments, facingpressure and suction sides of adjacent airfoils define a throat thatcorresponds to a gauge point, the throat defined as a minimum distancebetween the facing pressure and suction sides at a respective spanposition. The first passage axis of at least some passages of the firstset of crossover passages intersects the external wall at anintersection point. The intersection point is within 10% of a distancefrom the gauge point with respect to a length between the gauge pointand one of the leading edge and the trailing edge.

Although the different examples have the specific components shown inthe illustrations, embodiments of this disclosure are not limited tothose particular combinations. It is possible to use some of thecomponents or features from one of the examples in combination withfeatures or components from another one of the examples.

The various features and advantages of this invention will becomeapparent to those skilled in the art from the following detaileddescription of an embodiment. The drawings that accompany the detaileddescription can be briefly described as follows.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 schematically shows a gas turbine engine.

FIG. 2 schematically shows an airfoil arrangement for a section of a gasturbine engine.

FIG. 3 is a schematic view of airfoil span positions.

FIG. 4A illustrates a cross sectional view of adjacent airfoilsaccording to an example.

FIG. 4B illustrates a cross sectional view of adjacent airfoilsaccording to another example.

FIG. 5A illustrates a side view of a cooling arrangement with an airfoilshown in phantom.

FIG. 5B illustrates a cross sectional view of the cooling arrangement ofFIG. 5A along line 5B-5B.

FIG. 5C illustrates selected portions of the cooling arrangement of FIG.5B.

FIG. 6 illustrates an exemplary cooling arrangement according to anotherexample.

FIG. 7 illustrates another exemplary cooling arrangement.

FIG. 8 illustrates yet another exemplary cooling arrangement.

FIG. 9 illustrates a core for a cooling arrangement.

Like reference numbers and designations in the various drawings indicatelike elements.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle 15, while the compressor section 24drives air along a core flow path C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a first (or low) pressure compressor 44 and afirst (or low) pressure turbine 46. The inner shaft 40 is connected tothe fan 42 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48 to drivethe fan 42 at a lower speed than the low speed spool 30. The high speedspool 32 includes an outer shaft 50 that interconnects a second (orhigh) pressure compressor 52 and a second (or high) pressure turbine 54.A combustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 is arranged generally betweenthe high pressure turbine 54 and the low pressure turbine 46. Themid-turbine frame 57 further supports bearing systems 38 in the turbinesection 28. The inner shaft 40 and the outer shaft 50 are concentric androtate via bearing systems 38 about the engine central longitudinal axisA which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five (5:1). Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present invention isapplicable to other gas turbine engines including direct driveturbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption (‘TSFC’)”— is the industry standardparameter of 1 bm of fuel being burned divided by 1 bf of thrust theengine produces at that minimum point. “Low fan pressure ratio” is thepressure ratio across the fan blade alone, without a Fan Exit Guide Vane(“FEGV”) system. The low fan pressure ratio as disclosed hereinaccording to one non-limiting embodiment is less than about 1.45. “Lowcorrected fan tip speed” is the actual fan tip speed in ft/sec dividedby an industry standard temperature correction of [(Tram ° R)/(518.7°R)]^(0.5). The “Low corrected fan tip speed” as disclosed hereinaccording to one non-limiting embodiment is less than about 1150ft/second.

FIG. 2 shows selected portions of a section 53 of a gas turbine engine,such as a portion of the compressor section 24 or the turbine section 28of the engine 20 of FIG. 1. The section 53 includes a rotor 60 carryingone or more airfoils 61 for rotation about the central axis A. In thisdisclosure, like reference numerals designate like elements whereappropriate and reference numerals with the addition of one-hundred ormultiples thereof designate modified elements that are understood toincorporate the same features and benefits of the corresponding originalelements. In this example, each airfoil 61 includes a platform 62 and anairfoil section 65 extending in a spanwise or radial direction R fromthe platform 62 to a tip 64. The airfoil section 65 generally extends inan axial or chordwise direction C between a leading edge 66 to atrailing edge 68, and in a circumferential or thickness direction Tbetween suction and pressure sides S, P. A root section 67 of theairfoil 61 is mounted to the rotor 60, for example. It should beunderstood that the airfoil 61 can alternatively be integrally formedwith the rotor 60, which is sometimes referred to as an integrallybladed rotor (IBR). The airfoil 61 is rotatable about the central axisA.

A blade outer air seal (BOAS) 69 is spaced radially outward from the tip64 of the airfoil section 65. A vane 70 is positioned along the engineaxis A and adjacent to the airfoil 61. The vane 70 includes an airfoilsection 71 extending between an inner platform 72 and an outer platform73 to define a portion of the core flow path C. The section 53 includesmultiple airfoils 61, vanes 70, and blade outer air seals 69 arrangedcircumferentially about the engine axis A. Although the exemplarycooling arrangements discussed in the disclosure primarily refer to arotatable airfoil 61 or turbine blade, the teachings herein can also beutilized for another portion of the engine 20 including static airfoilssuch as vane 70 and other portions defining a gas path such as BOAS 69,for example. Also, although the teachings herein primarily refer to theleading edge or leading edge region, other portions of the airfoil canbenefit from the teachings herein, including trailing edge regions.

Referring to FIG. 3, span positions are schematically illustrated from0% to 100% in 10% increments to define a plurality of sections 74. Eachsection 74 at a given span position is provided by a conical cut thatcorresponds to the shape of segments of the core flow path C (FIG. 1),as shown by the large dashed lines. In the case of an airfoil 61 with anintegral platform, the 0% span position corresponds to the radiallyinnermost location where the airfoil 61 meets the fillet joining theairfoil section 65 to the platform 62. Span position may be relative tothe platform 62, such as 0% span at the platform 62 and 100% span at thetip 64, for example. In the case of an airfoil 61 without an integralplatform, the 0% span position corresponds to the radially innermostlocation where the discrete platform 62 meets the exterior surface ofthe airfoil section 65 (or the inner platform 72 of vane 70). A 100%span position corresponds to a section of the airfoil 61 at the tip 64(or the outer platform 73 of vane 70).

FIG. 4A shows an isolated view of a pair of adjacent airfoils designatedas leading airfoil 61A and following airfoil 61B at a first location,such as at a first span position or a location in the turbine section28. The airfoils 61A, 61B of FIG. 4A may correspond to high liftairfoils, for example, which are discussed in more detail below. Asshown, each airfoil 61A/61B is sectioned at a radial position betweenthe root section 67 and the tip 64 (FIGS. 2 and 3). A chord dimension(CD), which is the length between the leading and trailing edges66A/66B, 68A/68B of the airfoil 61A/61B, forms an angle, or staggerangle a, relative to the chordwise direction C or to a plane parallel tothe engine's central longitudinal axis A (FIGS. 2 and 3). The chorddimension (CD) may vary along the span of the airfoil 61A/61B.

Each airfoil 61A/61B can have an asymmetrical cross-sectional profilecharacterized by a mean camber line 75. The mean camber line 75 extendsbetween leading and trailing edges 66A/66B, 68A/68B to bisect athickness of the airfoil 61 in the circumferential or thicknessdirection T.

Each airfoil 61A/61B defines a camber angle (β) defined by a tangentialprojection of the mean camber line 75 at the leading and trailing edges66A/66B, 68A/68B. The camber angle (β) can differ at various spanpositions. For example, the camber angle (β) can decrease as spanposition increases. In some examples, the camber angle (β) is less thanabout 45 degrees, or more narrowly less than about 15 degrees, betweenthe 70% span position and the 100% span position, which may be utilizedwith high lift airfoil geometries, for example. For the purposes of thisdisclosure, the term “about” means ±3% of the stated value unlessotherwise disclosed.

The leading edges 66A, 66B or trailing edges 68A, 68B of the adjacentairfoils 61A, 61B are separated by a gap or circumferential pitch (CP)in the circumferential or thickness direction T to define an airfoilpassage 76. Facing pressure and suction sides P, S of adjacent airfoils61A, 61B define a throat 77. The throat 77 is defined as a minimumdistance between the facing pressure and suction sides P, S of adjacentairfoils 61A, 61B at a respective span position (see FIG. 3). The throat77 can include a component in the axial or chordwise direction C inaddition to a component in the circumferential or thickness direction T.The minimum distance corresponds to a gauge point GP along surfaces ofrespective facing pressure and suction sides P, S of adjacent airfoils61A, 61B. A location of the gauge point GP can vary depending on ageometry of the corresponding airfoils 61. In the illustrated example ofFIG. 4A, the gauge point GP is defined with respect to the suction sideS. FIG. 4B shows an isolated view of a pair of adjacent airfoilsdesignated as leading airfoil 61A′ and following airfoil 61B′ at asecond location, such as second span position, or at a location in thecompressor section 24, for example.

In some examples, the airfoil 61 is a “high lift” airfoil. For thepurposes of this disclosure, the term “high lift airfoil” means anairfoil design that has an airfoil shape that allows for reduced airfoilcount due to its ability to extract more work than a conventionalairfoil. High lift airfoils provide an improvement in efficiency andweight reduction. In using a high lift design, the airfoil stagnationpoint is shifted from a leading edge nose, where it is located on aconventional airfoil, to the pressure side towards the tip of theairfoil. In addition, the suction side gauge line, in which the gas Machnumber is at the greatest, on a high lift airfoil can occur much closerto the leading edge nose than a conventional airfoil. An exemplary highlift airfoil is disclosed in U.S. patent application Ser. No.14/767,768, entitled “Gas Turbine Engine High Lift Airfoil Cooling inStagnation Zone,” filed on Aug. 13, 2015 (now published as U.S.Publication No. 2016/0010463), the contents of which are incorporatedherein by reference in their entirety.

In examples, a cross section of the airfoil 61 at a span positionadjacent to the tip 64 is substantially thin and relatively long, andcan have a relatively large amount of twist about a radial axis of theairfoil 61. For example, a maximum thickness of the airfoil 61 at spanpositions between 70% span and 100% span can be less than 25% of amaximum thickness of the airfoil 61 at 0% span, or more narrowly lessthan 10% of the maximum thickness of the airfoil 61 at 0% span. In otherexamples, the chord dimension (CD) of the airfoil 61 at span positionsbetween 70% span and 100% span is greater than the chord dimension (CD)at 0% span, such as 125% or greater than the chord dimension (CD) at 0%span. Sections of example high lift airfoils are illustrated by airfoils161, 261, 361, 461 (FIGS. 5B-C and 6-8). A localized region of theairfoil 61 between the leading edge 66 and the gauge point GP, includingthe suction side S of high lift airfoils, can be susceptible to distressdue to exposure of relatively high pressures and/or temperatures alongthe core flow path C (FIG. 1), including oxidation, erosion, burnthroughand thermal-mechanical fatigue.

FIGS. 5A and 5B illustrate an exemplary cooling arrangement 180 for anairfoil 161, such as the airfoil 61 of FIG. 2. Although the exemplarycooling arrangements discussed in the disclosure primarily refer to acompressor or turbine blade, the teachings herein can also be utilizedfor another portion of the engine 20 such as vane 70 of FIG. 2, forexample.

At least one radial cooling passage 182 (only one shown for illustrativepurposes) is provided between pressure and suction sides P, S in thethickness direction T. Each radial cooling passage 182 extends from aroot section 167 through a platform 162 and toward a tip 164 tocommunicate coolant to various portions of the airfoil 161. Each radialcooling passage 182 is configured to receive coolant from a coolantsource CS (shown schematically). Exemplary coolant sources CS caninclude bleed air from an upstream stage of the compressor section 24,bypass air from the bypass flow path B (FIG. 1), or a secondary coolingsystem aboard the aircraft, for example.

The cooling arrangement 180 includes one or more feeding cavities 184and impingement cavities 186 extending in the radial direction R (onlyone feeding cavity 184 and one impingement cavity 186 shown in FIG. 5Afor illustrative purposes). One of the radial passages 182 or anothersource can communicate coolant to each of the feeding cavities 184 fordelivery to the impingement cavities 186.

The cavities 184, 186 can be formed in various locations of the airfoil161. In some examples, the impingement cavity 186 is bounded by anexternal wall 188 of the airfoil 161. The external wall 188 definespressure and suction sides P, S between leading and trailing edges 166,168 of the airfoil section 165. The airfoil section 165 can includemultiple feeding cavities 184 and/or impingement cavities 186 to providecooling to various portions of the airfoil section 165. For example, apair of impingement cavities 186 can be located at the leading edge 166,as shown in FIG. 5B. Impingement cavities 186 can be located at thetrailing edge 168 or between the leading and trailing edges 166, 168,also shown in FIG. 5B.

The airfoil 161 may include one or more film cooling holes or passages190 in fluid communication with the feeding cavities 184 and/orimpingement cavities 186 to provide film cooling to various surfaces ofthe airfoil 161. The film cooling passages 190 may extend from a surfaceof the impingement cavity 186, for example. In some examples, the filmcooling passages 190 are located within the external wall 188 at theleading edge 166, the trailing edge 168, or another location of theairfoil 161.

The airfoil 161 can include a thermal barrier coating (TBC) or coatingsystem 192 (shown in dashed lines in FIG. 5B) on an exterior surface ofthe airfoil section 165 to reduce heat transfer between the gas path andthe airfoil section 165. The thermal barrier coating 192 can include,but is not limited to, a ceramic material such as yttria-stabilizedzirconia (YSZ). The thermal barrier coating 192 can further include abond layer to facilitate adherence to the substrate.

One or more crossover passages 194 are located within internal wall 189of the airfoil section 165, and can be arranged to provide directionalor impingement cooling to relatively high heat load regions of theairfoil 161. The internal wall 189 includes one or more ribs 193 definedin a thickness of the airfoil that space apart or otherwise separate arespective feeding cavity 184 and impingement cavity 186 or multiples ofeither. Each rib 193 can define a reference plane RF extending in thespanwise or radial direction R along an adjacent surface of therespective impingement cavity 186. The reference plane RF can betransverse or perpendicular to mean camber line 175. The reference planeRF can be generally planar or can include a curvilinear component, forexample. The reference plane RF can also include an axial twist betweenplatform 162 and the tip 164.

Referring to FIG. 5C, with continued reference to FIGS. 5A and 5B, eachcrossover passage 194 extends between an inlet port 195 and an outletport 196 to connect a respective feeding cavity 184 and impingementcavity 186. In examples, each outlet port 196 of at least one, or morethan one, of the crossover passages 194 is defined at a span positionthat is between an 70% span position and the 100% span position (see,e.g., crossover passages 194A, 194B of FIG. 5A), and can be arrangedutilizing any of the techniques disclosed herein. Although only one setof crossover passages 194A, 194B is shown in FIG. 5, it should beappreciated that each of the impingement cavities 186A, 186B can beprovided with a respective set of crossover passages 194A, 194B eachincluding a plurality of crossover passages 194 at different spanpositions of the airfoil 161.

Each of the crossover passages 194 defines a passage axis PA thatintersects a surface of the impingement cavity 186 and the feedingcavity 184. The passage axis PA can be substantially linear or caninclude curvilinear portions between the inlet port 195 and outlet port196, and can intersect the external wall 188 at locations along aleading edge region 179 that defines the leading edge 166 of the airfoil161. For the purposes of this disclosure, the term “leading edge region”includes a region of the airfoil section 165 from the leading edge 166and within 20% of a distance along the external wall 188 between theleading and trailing edges 166, 168, unless otherwise disclosed.

The crossover passages 194 are arranged such that coolant provided tothe feeding cavity 184 is thereafter communicated to the impingementcavity 186. The coolant is communicated to the impingement cavity 186 toselectively provide impingement cooling to one or more portions of theexternal wall 188 of the airfoil 161. One or more of the crossoverpassages 194 can be arranged such that the respective outlet port 196 isspaced apart from the external wall 188 (e.g., crossover passage 194B).The crossover passages 194 are shown in FIG. 5A as having a uniformdistribution in the radial direction R. In other examples, the airfoil161 includes a non-uniform distribution of at least some of thecrossover passages 194 in the radial direction R. Although a particularquantity of crossover passages 194 is shown, the airfoil 161 can includefewer or more crossover passages 194.

Each passage axis PA defines a first or radial angle (γ) (FIG. 5A)relative to the reference plane RF. In some examples, the passage axisPA is substantially perpendicular to the reference plane RF, and theradial angle (γ) is approximately equal to 90 degrees such that thecrossover passage 194 is substantially parallel with the engine axis A.In other examples, the passage axis PA is transverse to the referenceplane RF, and radial angle (γ) is not equal to 90 degrees such that thecrossover passage 194 is transverse to the engine axis A and slopestowards one of the platform 162 or the tip 164.

The passage axis PA can define a second or passage angle (θ) (FIG. 5C)with respect to the reference plane RF such that the passage axis PA hasa component that extends in the chordwise direction C and such that thecrossover passage 194 ejects coolant onto surfaces of the impingementcavity 186 along the passage axis PA to cool a desired portion of theairfoil 161, such as portions of the external wall 188. The passage axisPA can include curvilinear portions between the inlet port 195 andoutlet port 196, for example, but the passage angle (θ) is defined by alinear projection of the passage axis PA from the intersection of thepassage axis PA and the reference plane RF at the outlet port 196.

The crossover passages 194 feeding a common one of the impingementcavities 186 can have the same passage angle (θ), or at least some oreach of the passage angles (θ) can differ to provide different coolingaugmentation to different portions of the external wall 188 bounding therespective impingement cavity 186. The crossover passage 194 can bedefined along various locations of the internal wall 189 or rib 193, andoutlet port 196 can be defined along various locations relative to thereference plane RF, including toward the suction side S or the pressureside P.

The airfoil 161 includes first and second (or a pair of) impingementcavities 186A, 186B that are bounded by external wall 188 at the leadingedge region 179, with no other impingement cavities being between theleading edge 166 and the cavities 186A, 186B. The cavities 186A, 186Bare separate and distinct cavities that provide dedicated impingementcooling to different portions of the airfoil 161 along the leading edgeregion 179, such as adjacent the respective suction and pressure sidesS, P. Feeding the impingement cavities 186A, 186B separately withcooling airflow can reduce a likelihood of the impingement cavity 186Aadjacent the suction side S being supplied with more cooling airflowthan necessary and a likelihood of the impingement cavity 186B adjacentthe pressure side P being supplied with insufficient cooling airflow dueto starvation, for example. The impingement cavities 186A, 186B canprovide relatively greater wetted surface area as compared to a singleimpingement cavity of the same volume, and can improve uniformity ofheat transfer between the cooling airflow and portions of the airfoil161 along the leading edge region 179.

Each impingement cavity 186, including cavities 186A, 186B, can extendsubstantially along a span of the airfoil section 165. In otherexamples, one or more impingement cavities 186′ can extend from the tip164 toward the root section 167, and one or more impingement cavities186″ can extend from the platform 162 or root section 167 toward the tip164, including any of the impingement cavities disclosed herein (one setof cavities 186′ and 186″ are shown in dashed line in FIG. 5A forillustrative purposes).

First and second feeding cavities 184A, 184B are separated from theimpingement cavities 186A, 186B by the ribs 193A, 193B of the internalwall 189. A pair of crossover passages 194A, 194B within ribs 193A, 193Bof the internal wall 189 connect the feeding cavities 184A, 184B withthe respective impingement cavities 186A, 186B to supply coolingairflow. In the illustrated example, the feeding cavities 184A, 184B areseparate and distinct cavities that each receive coolant from thecoolant source CS (FIG. 5A). The impingement cavities 186A, 186B are notdirectly connected to the radial cooling passage 182 (FIG. 5A) or thecoolant source CS and instead receive substantially all cooling airflowfrom the feeding cavities 184A, 184B. The crossover passages 194A, 194Bcan be sized to provide an adequate backflow margin to reduce alikelihood of ingestion of core air in core flow path C (FIG. 1) intoimpingement cavities 186A, 186B through cooling passages 190 or airflowfrom the impingement cavities 186A, 186B and into the feeding cavities184A, 184B.

The passage angles (θ_(A), θ_(B)) can be the same or can differ to ejectcooling airflow on discrete portions of the external wall 188. In theillustrated example of FIG. 5C, the passage angles (θ_(A), θ_(B)) aretransverse to respective reference planes RF_(A), RF_(B) and are angledtoward the suction and pressure sides S, P. Passage angle (θ_(A)) ofcrossover passage 194A can be defined such that the passage axis PA_(A)intersects a surface of impingement cavity 186A adjacent to the suctionside S. Passage angle (θ_(B)) of crossover passage 194B can be definedsuch that the passage axis PA_(B) intersects a surface of impingementcavity 186B adjacent to the pressure side P. In other examples, passageangle (θ_(A)) and/or passage angle (θ_(B)) is substantiallyperpendicular to the respective reference planes RF_(A), RF_(B) orangled inwardly toward the mean camber line 175.

The passage angle (θ) of at least some of the crossover passages 194 canchange in the radial direction R. In some examples, the passage angle(θ) of at least some or each of the crossover passages 194 changes inthe radial direction R such that the passage angles (θ) are generallyprogressively larger or smaller. In other examples, the crossoverpassages 194 are arranged such that the passage angles (θ) increase asspan position increases for each, or at least some of, the crossoverpassages 194. In yet other examples, the crossover passages 194 arearranged such that the passage angles (θ) decrease as span positionincreases for each, or at least some of, the crossover passages 194.

The passage axis PA of at least some of the crossover passages 194 canintersect the external wall 188 at a respective intersection point P1(FIG. 5C). The intersection point P1 may coincide with the respectivegauge point GP, or may be defined aft of gauge point GP′ (FIG. 5C). Saiddifferently, the intersection point P1 can be defined at a positionbetween the trailing edge 168 and the gauge point GP, inclusive of thegauge point GP. In some examples, the intersection point P1 is within±10% of a distance from the gauge point GP with respect to a lengthbetween the gauge point GP and one of the leading and trailing edges166, 168 along an external wall perimeter of the airfoil 161.

The passage axis PA of at least one, or each, of the crossover passages194 is arranged at a corresponding passage angle (θ) such that thepassage axis PA of the respective crossover passage 194 intersects theexternal wall 188 at a location between respective gauge point GP and anaftmost location of the impingement cavity 186 with respect to thechordwise direction C. In the illustrated example of FIG. 5C, theimpingement cavity 186 is free of any film cooling passages 190 aft ofthe gauge point GP. In alternative examples, the impingement cavity 186includes at least one film cooling passage 190′ (shown in dashed linesin FIG. 5C) aft of the gauge point GP and/or intersection point P1.

The arrangement of the impingement cavities 186 and crossover passages194 utilizing the teachings herein can provide dedicated coolingaugmentation to adjacent portions of the external wall 188 without theneed for introducing film cooling passages downstream of the gauge pointGP and/or intersection point P1 which can result in aerodynamic losses.

FIG. 6 illustrates a cooling arrangement 280 for an airfoil 261according to another example. Airfoil section 265 includes a pair ofimpingement cavities 286A, 286B defined in a leading edge region 279that are supplied with cooling airflow from a pair of feeding cavities284A, 284B. As shown, the airfoil 261 along the leading edge region 279is substantially symmetrical about mean camber line 275 such that theimpingement cavities 286A, 286B have substantially the same geometry andvolume and such that the feeding cavities 284A, 284B have substantiallythe same geometry and volume. A length of rib 291 that separatescavities 284A, 286A from cavities 284B, 286B follows the mean camberline 275. Crossover passages 294A, 294B that connect the impingementcavities 286A, 286B and the respective feeding cavities 284A, 284B arespaced apart from the mean camber line 275.

FIG. 7 illustrates a cooling arrangement 380 for an airfoil 361according to yet another example. Airfoil section 365 includes a pair ofimpingement cavities 386A, 386B defined in a leading edge region 379that are supplied with cooling airflow from a pair of feeding cavities384A, 384B. The impingement cavities 386A, 386B are arranged such thatmean camber line 375 intersects impingement cavity 386B and is spacedapart from impingement cavity 386A.

A rib 393 in the internal wall 389 separates the impingement cavities386A, 386B. A length of the rib 393 extends along a reference plane RRthat is substantially transverse to the mean camber line 375, with thereference plane RR intersecting at least one of the pressure and suctionsides P, S (e.g., suction side S) along the leading edge region 379. Therib 393 can provide reinforcement within a foreign object debris (FOD)region or impact area of the airfoil section 365. The FOD region can besusceptible to an increased risk of impact by FOD relative to otherportions of the airfoil 361. The reference plane RR intersects the meancamber line 375 to define a third or rib angle (β). In examples, the ribangle (β) is between 30 and 80 degrees, or more narrowly between 45 and60 degrees. The rib 393 can be defined such that inlet 395B and outlet396B of crossover passage 394B are forward of inlet 395A and outlet 396Aof crossover passage 394A with respect to the chordwise direction C.

Impingement cavity 386B defines a first volume. Impingement cavity 386Adefines a second, different volume. In some examples, the second volumeof impingement cavity 386A is less than about 75%, or more narrowly lessthan about 50% or half of the first volume of the impingement cavity386B. The different volumes can improve distribution of cooling airflowto the respective impingement cavities 386A, 386B based on coolingdemands adjacent the suction and pressure S, P sides of the airfoil 361.

FIG. 8 illustrates a cooling arrangement 480 for an airfoil 461according to another example. Airfoil section 465 includes a pair ofimpingement cavities 486A, 486B defined in a leading edge region 479that are supplied with cooling airflow from a common feeding cavity 484.The impingement cavities 486A, 486B are arranged such that mean camberline 475 intersects impingement cavity 486B and the common feedingcavity 484 but is spaced apart from impingement cavity 486A.

The common feeding cavity 484 can be formed from a stepped core todefine different regions 485A, 485B that are joined together at alocation adjacent the mean camber line 475 and that feed cooling airflowto respective inlets 495A, 495B, for example. In the illustratedexample, region 485A extends a different or lesser distance than region485B with respect to the chordwise direction C. The different regions485A, 485B provide for different positioning of the impingement cavities486A, 486B, including within leading edge region 479 that can have arelatively narrow thickness or profile in the thickness direction T.Both inlet 495B and outlet 496B of crossover passage 494B are forward ofinlet 495A and outlet 496A of crossover passage 494A with respect to thechordwise direction C. In other examples, outlet 496B but not inlet 495Bis forward of inlet 495A and outlet 496A.

FIG. 9 illustrates a portion of a core 597 according to an example. Thecore 597 can be formed or made via casting, additive, machining, or anymethod that can produce the geometric shape. The core 597 can beutilized in an additive manufacturing process and then used in a castingprocess, for example, to form the various portions of the coolingarrangements 180/280/380/480.

The core 597 includes a first portion 598A corresponding to the firstimpingement cavity 186A/286A/386A/486A, and a second portion 598Bcorresponding to the second impingement cavity 186B/286B/386B/486B, forexample. In examples, the first portion 598A defines a first volume, andthe second portion 598B defines a second, different volume including anyof the volumes discussed herein.

The core 597 includes a third portion 599A corresponding to the firstfeeding cavity 184A/284A/384A/484A, and a fourth portion 599Bcorresponding to the second feeding cavity 184B/284B/384B/484B, forexample. In embodiments, the third and fourth portions 599A, 599B areportions of a common portion 599′ (shown in dashed lines) thatcorresponds to common feeding cavity 484, for example.

Crossover connectors 500A, 500B, which may correspond to one or more ofthe crossover passages 194/294/394/494, couple or otherwise connect thefirst and third portions 598A, 599A and the second and fourth portions598B, 599B, respectively. The crossover connectors 500A, 500B can definerespective first and second sets of crossover passages194A/294A/394A/494A, 194B/294B/394B/494B at different span positions ofthe airfoil 161/261/361/461. The crossover connectors 500A, 500B can bearranged to define a passage angle (θ) according to any of thecorresponding crossover passages 194/294/394/494 disclosed herein.

The cooling arrangements 180/280/380/480 disclosed herein can improvecooling augmentation to selected portions of the airfoil or vane,including areas susceptible to distress, and can improve airfoildurability and component life. Airfoil designs including high liftairfoils can receive improved impingement cooling in the leading edgeregion, for example. The cooling arrangements 180/280/380/480 can alsoimprove efficiency of the engine 20 due to reducing cooling supplydemands, for example.

Although particular step sequences are shown, described, and claimed, itshould be understood that steps may be performed in any order, separatedor combined unless otherwise indicated and will still benefit from thepresent disclosure.

It should be understood that relative positional terms such as“forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like arewith reference to the normal operational attitude of the vehicle andshould not be considered otherwise limiting.

The foregoing description is exemplary rather than defined by thelimitations within. Various non-limiting embodiments are disclosedherein, however, one of ordinary skill in the art would recognize thatvarious modifications and variations in light of the above teachingswill fall within the scope of the appended claims. It is therefore to beunderstood that within the scope of the appended claims, the disclosuremay be practiced other than as specifically described. For that reasonthe appended claims should be studied to determine true scope andcontent.

What is claimed is:
 1. An airfoil for a gas turbine engine comprising:an airfoil section having an internal wall and an external wall, theexternal wall defining pressure and suction sides extending in achordwise direction between a leading edge and a trailing edge, whereinthe airfoil section defines a mean camber line extending between theleading and trailing edges to bisect a thickness of the airfoil section;a first impingement cavity and a second impingement cavity, both thefirst and second impingement cavities bounded by the external wall at aleading edge region that defines the leading edge, a first lineperpendicular to the mean camber line intersecting both the first andsecond impingement cavities, and the leading edge region being a regionof the airfoil section from the leading edge and within 20% of adistance along the external wall between the leading and trailing edges;a first feeding cavity separated from the first impingement cavity andfrom the second impingement cavity by the internal wall; a firstcrossover passage within the internal wall that connects the firstimpingement cavity and the first feeding cavity, wherein the firstcrossover passage defines a first passage axis that intersects a surfaceof the first impingement cavity; and a second crossover passage withinthe internal wall connected to the second impingement cavity, whereinthe second crossover passage defines a second passage axis thatintersects a surface of the second impingement cavity.
 2. The airfoil asrecited in claim 1, wherein the airfoil section extending in a spanwisedirection from a 0% span position to a 100% span position, the airfoilis a high lift airfoil, a camber angle is defined by a tangentialprojection of the mean camber line at the leading and trailing edges,and the camber angle is less than 15 degrees between a 70% span positionand the 100% span position.
 3. The airfoil as recited in claim 2,wherein the airfoil is a turbine blade.
 4. The airfoil as recited inclaim 2, wherein a rib separating the first impingement cavity and thesecond impingement cavity is transverse to the mean camber line.
 5. Theairfoil as recited in claim 2, wherein a length of a rib separating thefirst impingement cavity and the second impingement cavity follows themean camber line.
 6. The airfoil as recited in claim 5, wherein thesecond crossover passage connects the second impingement cavity and asecond feeding cavity, the first and second feeding cavity separated bythe length of the rib.
 7. An airfoil for a gas turbine enginecomprising: an airfoil section having an internal wall and an externalwall, the external wall defining pressure and suction sides extending ina chordwise direction between a leading edge and a trailing edge,wherein the airfoil section defines a mean camber line extending betweenthe leading and trailing edges to bisect a thickness of the airfoilsection; a first impingement cavity and a second impingement cavity,both the first and second impingement cavities bounded by the externalwall at a leading edge region that defines the leading edge, a firstline perpendicular to the mean camber line intersecting both the firstand second impingement cavities, and the leading edge region being aregion of the airfoil section from the leading edge and within 20% of adistance along the external wall between the leading and trailing edges;a first feeding cavity separated from the first impingement cavity andfrom the second impingement cavity by the internal wall; a firstcrossover passage within the internal wall that connects the firstimpingement cavity and the first feeding cavity, wherein the firstcrossover passage defines a first passage axis that intersects a surfaceof the first impingement cavity; a second crossover passage within theinternal wall connected to the second impingement cavity, wherein thesecond crossover passage defines a second passage axis that intersects asurface of the second impingement cavity; and wherein the first passageaxis intersects the external wall at an intersection point, theintersection point being between the trailing edge and a gauge pointdefined by the airfoil.
 8. The airfoil as recited in claim 7, whereinthe second crossover passage connects the second impingement cavity anda second feeding cavity, the second feeding cavity separated from thefirst impingement cavity and from the second impingement cavity by theinternal wall.
 9. The airfoil as recited in claim 7, wherein the firstpassage axis intersects the surface of the first impingement cavityadjacent to the pressure side, and the second passage axis intersectsthe surface of the second impingement cavity adjacent to the suctionside.
 10. The airfoil as recited in claim 7, wherein: the firstcrossover passage extends between a first inlet and a first outlet, thesecond crossover passage extends between a second inlet and a secondoutlet, and the first inlet is forward of the second outlet with respectto the chordwise direction; and the first outlet is forward of thesecond inlet with respect to the chordwise direction.
 11. The airfoil asrecited in claim 10, wherein the second crossover passage connects thesecond impingement cavity and the first feeding cavity.
 12. The airfoilas recited in claim 7, wherein the first impingement cavity defines afirst volume and the second impingement cavity defines a second volumethat is less than the first volume.
 13. A gas turbine engine comprising:an array of airfoils circumferentially distributed about an engine axis,each airfoil of the array of airfoils including an airfoil sectionhaving an internal wall and an external wall, the external wall definingpressure and suction sides extending in a chordwise direction between aleading edge and a trailing edge; and wherein the airfoil sectioncomprises: a first impingement cavity and a second impingement cavity,both the first and second impingement cavities bounded by the externalwall at a leading edge region that defines the leading edge; a firstfeeding cavity and a second feeding cavity separated from the firstimpingement cavity and from the second impingement cavity by theinternal wall; a first set of crossover passages within the internalwall that connect the first impingement cavity and the first feedingcavity, wherein each passage of the first set of crossover passagesdefines a first passage axis that intersects a surface of the firstimpingement cavity; and a second set of crossover passages within theinternal wall that connect the second impingement cavity and the secondfeeding cavity, wherein each passage of the second set of crossoverpassages defines a second passage axis that intersects a surface of thesecond impingement cavity; wherein facing pressure and suction sides ofadjacent airfoils define a throat that corresponds to a gauge point, thethroat defined as a minimum distance between the facing pressure andsuction sides at a respective span position; and wherein the firstpassage axis of at least some passages of the first set of crossoverpassages intersects the external wall at an intersection point, theintersection point being within 10% of a distance from the gauge pointwith respect to a length between the gauge point and one of the leadingedge and the trailing edge.
 14. The gas turbine engine as recited inclaim 13, comprising a compressor section and a turbine section, whereinthe array of airfoils are located in at least one of the compressorsection and the turbine section.
 15. The gas turbine engine as recitedin claim 13, wherein the airfoil section defines a mean camber lineextending between the leading and trailing edges to bisect a thicknessof the airfoil section, both the first and second impingement cavitiesare bounded by the external wall at the leading edge region, a firstline perpendicular to the mean camber line intersecting both the firstand second impingement cavities and the leading edge region being aregion of the airfoil section from the leading edge and within 20% of adistance along the external wall between the leading and trailing edges.16. The gas turbine engine as recited in claim 15, comprising acompressor section and a turbine section, wherein the array of airfoilsare located in at least one of the compressor section and the turbinesection.
 17. The gas turbine engine as recited in claim 16, wherein thearray of airfoils are rotatable blades in the turbine section.
 18. A gasturbine engine, comprising: an array of airfoils circumferentiallydistributed about an engine axis, each airfoil of the array of airfoilsincluding an airfoil section having an internal wall and an externalwall, the external wall defining pressure and suction sides extending ina chordwise direction between a leading edge and a trailing edge, andthe airfoil section defining a mean camber line extending between theleading and trailing edges to bisect a thickness of the airfoil section;and wherein the airfoil section comprises: a first impingement cavityand a second impingement cavity both bounded by the external wall at aleading edge region that defines the leading edge, a first lineperpendicular to the mean camber line intersecting both the first andsecond impingement cavities, and the leading edge region being a regionof the airfoil section from the leading edge and within 20% of adistance along the external wall between the leading and trailing edges;a first set of crossover passages within the internal wall that connectthe first impingement cavity and a first feeding cavity, wherein eachpassage of the first set of crossover passages defines a first passageaxis that intersects a surface of the first impingement cavity; and asecond set of crossover passages within the internal wall, wherein eachpassage of the second set of crossover passages defines a second passageaxis that intersects a surface of the second impingement cavity.
 19. Thegas turbine engine as recited in claim 18, wherein: facing pressure andsuction sides of adjacent airfoils define a throat that corresponds to agauge point, the throat defined as a minimum distance between the facingpressure and suction sides at a respective span position; and the firstpassage axis of at least some passages of the first set of crossoverpassages intersects the external wall at an intersection point, theintersection point being within 10% of a distance from the gauge pointwith respect to a length between the gauge point and one of the leadingedge and the trailing edge.
 20. The gas turbine engine as recited inclaim 19, wherein: the airfoil section includes a second feeding cavity,the first feeding cavity and the second feeding cavity separated fromthe first impingement cavity and from the second impingement cavity bythe internal wall; and the second set of crossover passages connect thesecond impingement cavity and the second feeding cavity.
 21. The gasturbine engine as recited in claim 19, wherein: each passage of thefirst set of crossover passages extends between a first inlet and afirst outlet, each passage of the second set of crossover passagesextends between a second inlet and a second outlet, and the first inletis forward of the second outlet with respect to the chordwise direction;the second set of crossover passages connect the second impingementcavity and the first feeding cavity, and the first outlet is forward ofthe second inlet with respect to the chordwise direction.